Summary The Bell 206L helicopter (registration C-FVIX, serial number45139) was in cruise flight at an altitude of about 700 feet above sea level when the pilot heard a sudden unusual noise and subsequently experienced an engine power loss. He lowered the collective and checked the instruments while scanning the area for a landing spot. The engine was still running; however, the turbine outlet temperature was climbing very rapidly and quickly exceeded the range of the gauge. The pilot subsequently raised the collective slowly, but the main rotor started to droop. He advised the two passengers of an engine failure and entered auto-rotation. While initiating a flared landing, he pulled the collective and confirmed no power from the engine as the low rotor horn sounded. The helicopter landed on a logging road near Tasu Creek, Queen Charlotte Islands, in the Sandspit area at 0829Pacific daylight time. The pilot shut down the engine immediately on landing. There were no injuries and no airframe damage. Ce rapport est galement disponible en franais. Other Factual Information The helicopter was equipped with a Rolls-Royce Allison 250-C20R engine, serial number CAE295208. At the site, a visual inspection revealed a large crack, about four inches long, in the lower weld joint of the ignitor and fuel nozzle mount at the back of the combustion section. The left-hand exhaust duct was also cracked at the weld on the forward side of the outer combustion case, about six inches down from the top. The duct also had impact damage which appeared to have originated from inside the engine. There was no visual damage from the substantial turbine outlet temperature (TOT)-indicated over-temperature that preceded the autorotation. There was a grinding noise heard from the engine when the blades were rotated backwards. Fuel and oil samples were retained; no obvious anomalies were noted. The lower chip detector was clean, and the oil did not contain obvious metallic particles. The upper chip detector and the freewheel chip detector were removed, and small amounts of metal were found. The compressor showed no signs of foreign object damage, but it could not be turned by hand. The engine had a total time since new (TTSN) of 6694.4hours. The times in service for the modular components are provided in AppendixA, with a summary of recent engine maintenance and repair. The technical records contained an entry on 05April2004, at 10817.0airframe hours, about 45engine hours before the occurrence, indicating an engine discrepancy related to a fuel-control unit adjustment due to hot starting. Another entry, dated 29November2003, indicated that the engine oil press/temp gauge was for a C20Bmodel engine and, therefore, had incorrect markings. Replacement parts were ordered, but there was no subsequent entry indicating that remedial action had been taken. Another entry, on 30September2003 at 10550airframe hours, indicated that the TOT harness was replaced due to a reading of 0.03ohms below minimum; the TOTgauge was also replaced due to it reading 15Clow. Following an engine teardown, several component parts were forwarded to the Transportation Safety Board Engineering Branch Laboratory for further examination and analysis. It was determined that the initiating event that led to the engine failure was the fracture of one blade on the second stage turbine wheel. The metallic debris from this fracture caused secondary engine damage as described below. For a more detailed account of the engine condition at examination, see LP068/04. A schematic of the turbine section of the engine is shown in AppendixB. A visual examination of the first-stage turbine wheel revealed many typeA and approximately four typeB cracks in the blade rim. Type A cracks are cracks in the platform, the surface between the blades. TypeB cracks are those that extend over the edge of the wheel rim and onto the face of the rim, but not more than 0.060inches. If cracking extends beyond0.060 inches into the rim, the wheel is rejected. It is considered normal to find cracks in the rim area. No cracks are allowed in the blades. The largest typeB crack observed was approximately 0.030inches on the leading-edge rim face, and no typeB cracks were observed on trailing edge rim face. Rub damage was observed on several blade tips covering an arc of approximately 180. Likewise, rub damage was observed on the inner stage seal, covering an arc of 180 as well. Some fretting wear was noted on the curvic coupling. Examination of the second-stage turbine wheel revealed one failed blade (see Photo1). Records indicate that the second-stage turbine wheel had accumulated 375.8hours TTSN, out of an allowable service life of 1775hours. An optical examination indicated blade failure as a result of fatigue cracking, with multiple initiation sites in the fillet area on the convex side of the blade. Fatigue cracking accounted for approximately 75percent of the fracture, with the remainder being an instantaneous overstress rupture. While the majority of the fracture face was normal to the blade axis, the origin area was almost parallel to the blade axis or radial to the turbine wheel. The blade following the failed blade had been bent opposite to the direction of rotation. Bending of this blade effectively reduced its overall length, preserving the pre-event tip condition. Further examination of the bent blade revealed minimal tip rub, and no blueing or mushrooming. Significant tip rub was observed on 75percent of the blades, from the failed blade over an arc of approximately 270clockwise from a leading-edge view. Rubbing on these blades had produced smearing damage to the tips of the blades with blued and mushroomed material at the outer diameter on the cambered face. The balance piston seal (large labyrinth seal) also showed rub damage around approximately 270of arc, consistent with the blade tip rub. Additionally, rub damage was noted on approximately 80percent, or 300of arc, of the inner stage seal, again consistent with the blade tip rub. Secondary smearing damage was observed on the trailing edge of several blades, consistent with interaction with the liberated piece of failed blade. The first- and second-stage wheels were mated at the curvic coupling, according to the reference marks vibro-etched on the hubs. Comparison of the blade tip rub showed both wheels were affected in the same general area, with the worst rub damage approximately 90ahead of the failed blade location. A scanning electron microscope examination of the second-stage turbine wheel blade fracture revealed multiple adjacent origin areas. Examination of the surface of the crack showed an oxidation layer that served to delineate a clear underlying pattern of fatigue striations. The fracture topography at the initiation site was consistent with low-cycle fatigue cracking as the initial mode of failure, with the observation of beach marks and widely spaced striations. The low-cycle fatigue cracking progressed radially inward toward the hub and then turned normal to the blade axis, progressing in a high-cycle mode. The blade eventually failed in the overstress extension of the high-cycle fatigue cracking. Secondary cracking was also observed in the fillet area, below the fracture plane. Energy dispersive x-ray analysis indicated the blade material was consistent semi-quantitatively with IN713C, the specified material. A transverse section was taken through the fracture origin of the second stage turbine wheel blade and mounted for metallurgical analysis. Two secondary cracks were observed in the fillet below the fracture. Chemical etching revealed a typical micro-structure with a visible gamma prime phase and well-defined grain boundaries. Additional sections taken at adjacent blade roots showed similar cracks in the fillet radius. The uniform distribution of the gamma prime phase indicated that there had been no overheating. Direct Rockwell hardness testing averaged 40Rockwell"C" (HRC) (equivalent ultimate tensile strength of 182Kpsi), which is within the manufacturer's specified maximum of 42HRC. No metallurgical anomalies were observed. Dimensional analysis of the second-stage nozzle, which had secondary damage, was completed at the manufacturer's facility in Indianapolis, Indiana. The secondary damage to the nozzle primarily affected the outer edge; therefore, the results were still considered meaningful. The analysis produced a blade path diameter that corresponds to a second-stage turbine blade tip clearance of 0.012inches, within the specification of 0.010to 0.016inches. The analysis also showed that, except for the damaged areas, the second stage blade path of the nozzle was within the specification for roundness. A blade from the first-stage turbine wheel was sectioned longitudinally to examine the gamma prime phase of the micro-structure for signs of overheating. The gamma prime phase was uniformly distributed throughout the blade, suggesting that long-term overheating did not occur. The third-stage turbine wheel showed significant leading-edge impact damage. The trailing edge was unremarkable and the curvic coupling was clean with minimum fretting wear. Minimum wear was observed on the inner stage seal. The power turbine support showed that both the inner and outer balance piston seals had significant rub damage around the entire circumference. The third-stage nozzle shield was deformed as a result of impact damage and could not be easily removed. The third-stage nozzle shield saddle had a major through-thickness gouge, at the nine o'clock position from a leading-edge view, consistent with impact from the failed second-stage blade. The outer combustion case showed a four-inch gaping crack along the boss weld below the fuel nozzle and ignitor ports. The fracture exhibited features consistent with fresh, low-cycle fatigue, considered to be the result of engine vibration after the second-stage turbine wheel blade failure. The Rolls Royce 250-C20R Series Operation and Maintenance Manual, Table4 - Measured Gas Temperature Limits (TOT) in AppendixC, provides actions to be taken when measured gas turbine outlet temperature (TOT) limits are exceeded during starting and shutdown and/or during power transients. About 45engine hours before the occurrence, the operator documented that the engine fuel-control unit was adjusted because of hot starts, but the amount by which the temperature was exceeded was not recorded. A turbine special inspection is recommended when TOT limits are exceeded. Reportedly, the TOT would briefly peak in the range of 810C to 830C. Hot starting events are not recorded by this helicopter's instrumentation. During these "hot starts," apparently the temperature/time limits that trigger a turbine inspection were never reached, after making allowances for the 0.03ohms (0.6C) TOT harness discrepancy and a 15C error from the incorrect TOT gauge (replaced at 63.1hours). Apparently, the ultimate temperature (927C) for which a turbine inspection is recommended was never reached during the so-called hot starts. The TOT harness may have had a 0.03ohms discrepancy, which would equate to 0.6C, and the incorrect TOT gauge was replaced at 63.1hours due to a 15C error. Examination of the rear support and diffuser ring revealed that approximately 25percent of the major diameter seal was missing. No corrosion was observed and the coloration indicated that the section of seal may have been gone for some time. Further examination showed that the seal was dis-bonding from the substrate. Dis-bonding was observed at both ends of the seal fracture, consistent with bond failure. Records indicated that the seal was installed about 700airframe hours prior to the occurrence. The fractured pieces of the soft major diameter seal material would most likely be broken down to fine particles and expelled safely through the rear support vent. A slight loss in engine efficiency would be the expected outcome. Failure of this seal was not considered contributory to the engine failure.